1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to composite turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an aero engine to power an aircraft or an industrial gas turbine engine to produce electric power, a hot gas flow is passed through a turbine to extract mechanical power. The efficiency of the engine can be increased by passing a higher temperature gas flow through the turbine. However, the highest temperature allowed is dependent upon the material properties of the turbine components, especially the first stage stator vanes and rotor blades. Thus, complex internal cooling circuits have been used to provide convection, impingement and film cooling to the airfoils to allow for higher temperatures.
Airfoils made entirely of a ceramic material has been proposed because the ceramic material can be exposed to a much higher temperature than the modern super-alloys. However, ceramic matrix composites (CMC) are brittle and do not have the strength and rigidity to operate within the engine. Some prior art turbine blades and vanes have been proposed in which a ceramic material is secured to a metallic material to form a composite turbine airfoil. However, the thermal coefficients of the ceramic and metallic materials are so different that high stresses are developed that also lead to problems for use in the engine.
Another design problem with turbine blades is providing for cooling air. The pressurized cooling air use to cool the blade typically is bled off from the compressor of the engine. The work to compressor the cooling air is not used to produce power in the engine, and is therefore wasted. Providing for a turbine blade that would require less cooling air would increase the efficiency of the engine.
Metallic turbine blades that are exposed to the high temperature of the turbine also suffer from thermal metallic fatigue (TMF) which tends to reduce the useful life of the blade. Hot spots that occur around the blade due to lack of adequate cooling also can cause erosion problems that will shorten the blade life.
For the industrial gas turbine engine design, large turbine blades can be heavy due to the requirement that the blade be made with enough strength to withstand the centrifugal loads from rotation under high temperatures that cause problems with creep. Turbine blade design attempts to produce the lightest weight blade that will withstand the high temperature and mechanical loads from engine operation. A lightweight blade design would allow for a higher AN2 which leads to higher engine efficiency.
It is therefore an object of the present invention to provide for a turbine blade with a reduced blade cooling flow requirement over the prior art turbine blades.
It is another object of the present invention to provide for a turbine blade with a reduced hot gas side convection surface required to be cooled for the spar.
It is another object of the present invention to provide for a turbine blade which eliminates the TMF issue normally experienced in the near wall cooling design of a turbine blade.
It is another object of the present invention to provide for a turbine blade with a high temperature resistant material on the pressure side of the blade which will eliminate the main body and the pressure side tip edge film cooling required in the prior art blades and thus reduce the blade total cooling flow demand and reduce the manufacturing complexity.
It is another object of the present invention to provide for a turbine blade with a light weight construction to allow for the turbine to be designed with a higher AN2 that the prior art turbine blades.